Spacecraft inertial attitude and rate sensor control system

ABSTRACT

Disclosed is a method and apparatus (1&#39;; 1&#34;) for tracking a stellar body (22) using a telescope (9; 32) of a spacecraft (e.g., a satellite) (10; 10&#34;). In accordance with an embodiment of the invention, the telescope (9; 32) is provided with gimbal supports (18a; 18b), and is maneuverable relative to the spacecraft (10; 10&#39;). The stellar body (22) is acquired by the telescope (9; 32) so that the stellar body (22) is within the field of view (FOV) of the telescope (9; 32). After the stellar body (22) is acquired, an operation is performed for controlling the attitude of the spacecraft (10; 10&#39;) to within pre-established deadband limits, and, as a result, the spacecraft (10; 10&#39;) and telescope (9; 32) are each assumed to have a desired orientation relative to the stellar body (22). The stellar body tracking method of the invention is performed so as to maintain the telescope (9; 32) in an orientation wherein the stellar body (22) is within the field of view (FOV) of the telescope (9; 32), even if it occurs that one or more environmental disturbance forces impinge on the spacecraft (10; 10&#39;) and cause an undesired variation in the orientation of the spacecraft (10; 10&#39;) relative to the stellar body (22). The stellar body tracking method of the invention controllably points a pointing direction (line of sight) of the stellar body tracking system (1&#39;; 1&#34;) independently of the spacecraft (10; 10&#39;), for tracking the stellar body (22). In accordance with an aspect of this invention, spacecraft inertial position and rate information is derived for use in controlling the attitude of the spacecraft (10; 10&#39;). In accordance with an embodiment of the invention, the pointing direction of the telescope (9; 32) is controllably pointed by maneuvering at least one of the telescope (9; 32) and a mirror (13&#34;) of the telescope (9; 32).

FIELD OF THE INVENTION

This invention relates generally to spacecraft attitude measurementsystems and, in particular, to a spacecraft inertial attitude and ratesensor control system.

BACKGROUND OF THE INVENTION

While a spacecraft is in an orbit, the spacecraft may be subject tovarious external disturbance forces which can produce a moment about thespacecraft's center of mass, causing the spacecraft's attitude tochange. This change in the spacecraft's attitude is undesirable in thatit can result in the spacecraft's payload not being correctly orientedto a selected location. As such, conventional spacecraft often includeattitude control systems which enable the attitude of the spacecraft tobe controlled within pre-established deadband limits. Such attitudecontrol systems often operate by employing conventional mechanisms suchas, e.g., gyros (e.g., laser gyros) and/or conventional star trackers,to enable spacecraft attitude-related information to be detected for usein controlling the spacecraft's attitude.

The use of such conventional mechanisms for detecting spacecraftattitude-related information has a number of associated drawbacks. Byexample, conventional gyros that are employed for detecting spacecraftinertial rate information are often expensive. Also, in conventionalstar trackers, which typically employ a large field of view (e.g.,8°×8°) telescope having a Charge Coupled Device (CCD), the CCDs and theelectronic circuits (e.g., such as those needed for providing clockdrive signals and video signal processing) associated with the CCDs tendto be expensive, and the electronic circuits are often complex.Moreover, the performance of such components can become degraded if thecomponents are exposed to solar radiation (the radiation hardness of atypical CCD is limited), and thus appropriate shielding must often beprovided for protecting the components from exposure to solar radiation.

In view of these considerations, it can be appreciated that it would bedesirable to provide a novel and inexpensive apparatus and technique forenabling spacecraft attitude-related information to be detected, forsubsequent use in controlling the spacecraft's attitude, and whichovercomes the above-described drawbacks associated with the prior art.

OBJECTS AND ADVANTAGES OF THE INVENTION

It is a first object and advantage of this invention to provide animproved apparatus and technique for providing spacecraft inertialposition information, for subsequent use in controlling the attitude ofthe spacecraft.

It is a second object and advantage of this invention to provide animproved apparatus and technique for providing spacecraft inertial rateinformation, for subsequent use in controlling the attitude of thespacecraft.

It is a further object and advantage of this invention to provide agimballed telescope for a spacecraft, and a technique for controllingthe position of the telescope for enabling the telescope to track apredetermined stellar body.

SUMMARY OF THE INVENTION

The foregoing and other problems are overcome and the objects of theinvention are realized by a method and apparatus for tracking a stellarbody using a telescope of a spacecraft, in accordance with thisinvention. In accordance with the invention, the telescope is providedwith gimbal supports, and may be maneuvered (i.e., rotated about atleast two axes) relative to the spacecraft.

It is assumed that the stellar body is acquired by the telescope (thestellar body is within the field of view of the telescope), and thatsubsequently an operation is performed for controlling the attitude ofthe spacecraft to within pre-established deadband limits. As a result ofthis operation, the spacecraft and telescope are each assumed to have adesired orientation relative to the stellar body.

The stellar body tracking method of the invention controls the positionof the telescope to enable the telescope to maintain an orientationwherein a desired stellar body is within the field of view of thetelescope, even if it occurs that one or more environmental (e.g.,solar) forces impinge on the spacecraft and cause an undesired variationin the orientation of the spacecraft relative to the stellar body. Thestellar body tracking method in accordance with this invention comprisessteps of a) detecting an undesired motion of the spacecraft which altersthe orientation of the spacecraft relative to the stellar body; and b)maneuvering the telescope during the undesired motion of the spacecraftso that the telescope is maintained in an orientation wherein thestellar body is within the field of view of the telescope. Themaneuvering step controllably points a pointing direction (line ofsight) of the telescope independently of the spacecraft, for enablingthe stellar body to be tracked.

In accordance with another aspect of this invention, spacecraft inertialposition and rate information is derived for use in controlling theattitude of the spacecraft. By example, in accordance with this aspectof the invention steps are performed of detecting a rate of motion ofthe telescope about at least one axis during the maneuvering of thetelescope, detecting a position of the telescope during the maneuveringof the telescope, and determining at least one of a position of thespacecraft and a rate of the undesired motion of the spacecraft, basedon the detected position of the telescope and the detected rate ofmovement of the telescope about the at least one axis.

Further in accordance with this invention, a step is performed ofproviding at least one torque to the spacecraft for controlling theattitude of the spacecraft within pre-established deadband limits, basedon the determined position of the spacecraft and the determined rate ofthe undesired motion of the spacecraft.

In accordance with another embodiment of the invention, the telescope isrotatable about a single axis, and includes a rotatable mirror that mayalso be maneuvered during the undesired motion of the spacecraft formaintaining the stellar body within the field of view of the telescope.

BRIEF DESCRIPTION OF THE DRAWINGS

The above set forth and other features of the invention are made moreapparent in the ensuing Detailed Description of the Invention when readin conjunction with the attached Drawings, wherein:

FIG. 1 depicts a Spacecraft Inertial Attitude and Rate Sensor ControlSystem (SIARSCS) that is constructed in accordance with an embodiment ofthis invention, wherein a gimballed telescope of the SIARSCS isrepresented in a perspective view.

FIGS. 2a and 2b show a representation of exemplary angular positionvalues and angular velocity values of the gimballed telescope of FIG. 1,as detected by respective position and rate sensor/encoder blocks of theSIARSCS of FIG. 1 over an exemplary time period, and FIG. 2c shows arepresentation of telescope inertial position values interpolated by acontroller of the SIARSCS of FIG. 1, superimposed over the exemplaryvalues of FIG. 2a.

FIG. 3a-3c show various orientations of a spacecraft and the telescopeof FIG. 1 relative to a stellar body, as viewed from a perspectivelooking down on the spacecraft, the telescope, and the stellar body.

FIGS. 4a-4c show a radiation sensitive surface 16a of a quad celldetector of the telescope of FIG. 1a, and also shows a representation ofstar light energy 21 incident on the surface 16a of the quad cell forthe various orientations of the spacecraft and telescope depicted inFIGS. 3a-3c, respectively.

FIGS. 5a and 5b depict a method in accordance with this invention forcontrolling the position of the telescope of FIG. 1.

FIG. 6 depicts the spacecraft of FIGS. 3a-3c, the telescope of thespacecraft, and electronic components 26 of the SIARSCS of FIG. 1,wherein the structure of the spacecraft prevents solar radiation 12impinging on the spacecraft from coming into contact with the electroniccomponents 26, in accordance with an aspect of this invention.

FIG. 7 depicts a Spacecraft Inertial Attitude and Rate Sensor ControlSystem (SIARSCS) that is constructed in accordance with anotherembodiment of this invention.

FIG. 8 illustrates a closed-loop control system for controlling theposition of the telescope of FIG. 1 in accordance with this invention.

Identically labeled elements appearing in different ones of the figuresrefer to the same elements but may not be referenced in the descriptionfor all figures.

DETAILED DESCRIPTION OF THE INVENTION

FIG. 1 illustrates a block diagram of a Spacecraft Inertial Attitude andRate Sensor Control System (SIARSCS) 1' that is constructed inaccordance with an embodiment of this invention. The SIARSCS 1'comprises a telescope 9, a telescope drive mechanism (also referred toas a telescope drive) 5, a position sensor/encoder block 3a, a ratesensor/encoder block 3b, a controller 1, a spacecraft attitudedetermination and control system (SADACS) 8, and attitude adjustmentactuators 11, which may include, for example, thrusters, momentumwheels, and/or magnetic torquers. Sensors 12, which may include, forexample, gyros, an earth sensor, and/or a sun sensor, may also beprovided for taking measurements that are used in providing attitudecontrol, although it is not necessary that these elements be employed,as will be described below.

Bearings 18a are disposed between a surface 9a of the telescope 9 and astructural portion 19 of a spacecraft 10, and provide gimbal supportsfor the telescope 9 to allow the telescope 9 to be rotated about axes x"and z". The telescope drive mechanism 5 is drivably-engaged with thesurface 9a of the telescope 9. The telescope drive mechanism 5 iscontrollable by the controller 1 for being actuated to provide selectedtorques to the telescope 9 for causing the telescope 9 to rotate aboutone or more of the axes x" and z" by selected angular displacementsrelative to the spacecraft structure 19. Preferably, the telescope drivemechanism 5 and the bearings 18a have a collective capability forenabling the telescope 9 to be rotated about the axis z" by at least 50degrees and about the x" axis by at least ±4 degrees, thereby allowingthe telescope 9 to scan over ±4 degrees in elevation (±8 degrees line ofsight) and over ±50 degrees in azimuth. Also, in a preferred embodimentof this invention, the telescope 9 has a field of view (FOV) of about±0.5 degrees.

It should be noted that any suitable type of drive mechanism may beemployed for the telescope drive 5, including, by example, anelectro-mechanical or inertial drive mechanism.

In the preferred embodiment of the invention, the telescope 9 comprisesa mirror 13, a primary lens 14, a secondary lens 15, a quad celldetector (also referred to as a quad cell or light-sensitive optic) 16,and a centroid error detection block (CEDB) 17, which operates in amanner as will be described below. Wavelengths 2 (such as those emittedby a stellar body within the field of view (FOV) of telescope 9)received by the telescope 9 through an aperture 9b of the telescope 9are reflected from the mirror 13 to the primary lens 14, which thenfocusses the wavelengths it receives to the secondary lens 15. Thoseones of the wavelengths that are reflected to the secondary lens 15 arefocussed by the secondary lens 15 to a radiation sensitive surface 16aof the quad cell 16. A representation of star light energy 21 incidenton the surface 16a of the quad cell 16 is shown in FIG. 1.

The quad cell 16 is responsive to the star light energy 21 incident onthe surface 16a of the quad cell 16 for producing signals correspondingto the star light energy 21. The quad cell 16 is controllable by thecontroller 1 for reading out these signals to the CEDB 17. Based onsignals received by the CEDB 17 from the quad cell 16, the CEDB 17computes the centroid of the distribution of star light energy 21 on thequad cell surface 16a. The CEDB 17 also calculates the position andoffset of the centroid relative to an x'-y' coordinate system defined bythe quad cell 16, and outputs information to the controller 1representing the calculated position and offset of the centroid relativeto the x'-y' coordinate system, as will be further described below.

It should be noted that any suitable known type of small area detectorfor maintaining fine lock may be employed for the quad cell 16,including, by example, a small area CCD device. Preferably, the quadcell 16 has a field of view of ±0.5 mrad. Also, any suitable type ofdevice/circuit and technique may be employed for computing, based on thequad cell outputs, the centroid of distribution of star light energy 21on the quad cell surface 16a and the position and offset of the centroidrelative to the coordinate system defined by the quad cell 16. Moreover,it is within the scope of this invention to employ other types ofdetector arrays besides quad cell detectors, so long as a position of anenergy distribution incident on a radiation sensitive surface of thedetector array can be determined relative to predetermined referencecoordinates/axes.

The position sensor/encoder block 3a and the rate sensor/encoder block3b will now be described. The position sensor/encoder block 3a (whichmay include, for example, one or more shaft position indicators andencoders) detects the angular position (i.e. , the arc of rotation) ofthe telescope 9 relative to individual ones of the axes z" and x", andtranslates the detected angular position to corresponding information(e.g., a series of digital pulses) which represents the detected angularposition of the telescope 9. This information is provided to thecontroller 1, which employs the information in a manner as will bedescribed below.

For cases where the telescope 9 is rotated about one or more of the axesx" and z" (as will be described below), the rate sensor/encoder block 3bdetects the angular velocity of the telescope's rotation about theseaxes x" and/or z", and translates the detected angular velocity abouteach axis z" and x" to corresponding information (e.g., a series ofdigital pulses). The information, which represents the detected angularvelocity of the telescope 9 about the individual axes z" and x" isoutput by the rate sensor/encoder block 3b to the controller 1, whereinit is employed in a manner as will be described below. It should benoted that it is within the scope of this invention to calculate withinthe controller 1 the angular velocity of the telescope 9 about eachindividual axis z" and x" based on outputs of the positionsensor/encoder block 3a, and thus, in an alternate embodiment of theinvention a separate rate sensor/encoder block 3b need not be employed.

The spacecraft attitude determination and control system (SADACS) 8 isassumed to have a capability for controlling the attitude of thespacecraft 10 for maintaining the attitude of the spacecraft 10 withinpre-established deadband limits, based on measurements made by theblocks 3a and 3b and/or the sensors 12. In accordance with an aspect ofthe invention, however, and as was previously mentioned, the sensors 12need not be employed for providing such measurements, and the spacecraftattitude may be controlled based solely on measurements made by theblocks 3a and 3b. This aspect of the invention will be further explainedbelow.

The technique employed by the SADACS 8 for controlling the attitude ofthe spacecraft 10 may be any suitable type of technique for providingspacecraft attitude control. By example, reference may be had to U.S.Pat. No. 5,348,255, issued Sep. 20, 1994, entitled System and Method forSensing Attitude of a Spacecraft with Equilized Star Tracker ErrorsAlong Three Orthogonal Axes, by Rene Abreu, wherein an attitude controlprocessor is described which may be one suitable embodiment of theSADACS 8. The SADACS 8 controls the spacecraft attitude by activatingselected ones of the spacecraft attitude adjustment actuators 11, suchas thrusters, to provide appropriate roll, pitch, and/or yaw torques forcontrolling the spacecraft attitude.

Before describing the method of the invention in detail, somepreliminary considerations will first be made. For the purposes of thefollowing description, it is assumed that the location of the telescope9 on the spacecraft 10 is such that the axis z" (FIG. 1) passing throughthe telescope 9 is the same as axis z (yaw) represented in FIGS. 3a-3c,which show various orientations of the spacecraft (e.g., a satellite)10, telescope 9, and the field of view (FOV) and line of sight (LOS)(also referred to as a pointing direction) of the telescope 9, relativeto a stellar body (e.g., assumed to be a star) 22, from a perspectivelooking down on these elements 10, 9, FOV, LOS, and 22. Preferably, thespacecraft 10 is in a low earth orbit (LEO), although other orbitalconfigurations may be employed.

FIG. 3a is an exemplary representation of the spacecraft 10 and the star22 at a first instance in time. It is assumed that at this time, thestar 22 (e.g., which may have a brightness of approximately 5 Mv) isacquired by the spacecraft telescope 9. The technique employed by thespacecraft 10 for acquiring the star 22 may be any suitable known typeof stellar body acquisition technique. For example, the acquisitiontechnique may employ information from a star catalog in an 8 degree×10degree Field of Regard of the gimballed telescope 9. Preferably, twofocal planes are employed for reducing acquisition time. For example, inone embodiment one of the focal planes is provided by employing sixteencourse acquisition detectors (a large field of view) in the telescope 9,and the other focal plane is provided by the quad cell 16, whichprovides fine lock tracking.

It is also assumed that after the star 22 is acquired, the spacecraftattitude determination and control system 8 operates to control theattitude of the spacecraft 10 so as cause the spacecraft attitude to bewithin pre-established deadband limits.

The orientation of the spacecraft 10, telescope 9, and the telescope FOVand LOS relative to the star 22, as depicted in FIG. 3a, is assumed tobe a desired orientation for the spacecraft 10, telescope 9, FOV and LOSrelative to the star 22. As is represented in FIG. 3a, the star 22 isassumed to be located within the field of view (FOV) and line of sight(LOS) of the telescope 9. Also, it is assumed that while the spacecraft10, telescope 9, FOV, and LOS are oriented relative to the star 22 inthe manner shown in FIG. 3a, a centroid of a distribution of star lightenergy 21 (resulting from light received by telescope 9 from star 22)incident on the quad cell surface 16a is located at an intersection ofaxes x' and y' defined by the quad cell 16, as is represented in FIG.4a. This is assumed to be an expected position of the centroid of thestar light energy 21 on the quad cell 16. For the purposes of thisdescription, this expected position of the centroid of the star lightenergy 21 (i.e., the intersection of axes x' and y') is hereinafter alsoreferred to as a line of sight reference point.

As was previously described, while a spacecraft, such as the spacecraft10, is in an orbit, various forces (e.g., solar forces and/or otherenvironmental forces) may impinge on the spacecraft 10 and cause achange in the attitude of the spacecraft 10. As can be appreciated, andassuming that the telescope drive 5 is not activated so that thetelescope 9 is stabilized, a change in the attitude of the spacecraft 10also can cause a change in the orientation of the telescope 9 and thetelescope's field of view (FOV) and line of sight (LOS) relative to thestar 22. This can cause the centroid of the star light energydistribution 21 on the quad cell surface 16a to "drift" (i.e., displace)away from the line of sight reference point by a displacementcorresponding to the amount of displacement of the telescope 9 (LOS, andFOV) from the position shown in FIG. 3a. In accordance with an aspect ofthis invention, the SIARSCS 1' operates as a closed loop control systemfor performing a fine tracking technique to maintain the centroid of thedistribution of star light energy 21 substantially at the line of sightreference point defined by quad cell 16. Also, the technique of theinvention enables the spacecraft's inertial rate and position to beestimated so that the attitude of the spacecraft can be controlled basedon these estimates. The manner in which the SIARSCS 1' operates will nowbe described in detail.

It is assumed that, as a result of, e.g., solar and/or otherenvironmental forces impinging on the spacecraft 10, the spacecraft 10is caused to yaw at a constant angular velocity such that theorientation (attitude) of the spacecraft 10, and hence, the orientationof the telescope 9, LOS, and FOV, relative to the star 22, variesaccordingly. This occurrence is also represented by block (A) of theflow diagram of FIG. 5a. A representation of the orientation of elements9, 10, LOS, and FOV relative to star 22 at a second instance in timethat occurs during the spacecraft yaw rotation is shown in FIG. 3b. Ascan be seen in FIG. 3b, at this second instance of time, the spacecraft10 has been displaced about axis z by an angle of (θ₁).

It is also assumed that, as a result of the change in the attitude ofthe spacecraft 9, and the corresponding change in the orientation of thetelescope 9, LOS, and FOV relative to the star 22, the centroid of thedistribution of star light energy 21 on the quad cell surface 16a isdisplaced from the line of sight reference point by an angulardisplacement corresponding to the spacecraft's angular displacement(θ₁). By example, a representation of the location of the star lightenergy distribution 21 on the quad cell surface 16a at the secondinstance of time is represented in FIG. 4b. The amount of displacementof the star light energy distribution 21 from the line of sightreference point is corresponds to the angle (θ₁) of displacement of thespacecraft 10. For the purposes of this description, this displacementof the star light energy distribution 21 from the line of sightreference point is assumed to be a maximum acceptable displacement ofthe star light energy distribution 21 from the line of sight referencepoint.

Assuming that the quad cell 16 is controlled by the controller 1 forreading out signals corresponding to the star light energy 21 incidenton the quad cell surface 16a at the second instance of time, and that,as a result, the signals are forwarded from the quad cell 16 to the CEDB17, the CEDB 17 computes the centroid of the distribution of star lightenergy 21 on the quad cell surface 16a, and calculates the position ofthe centroid on the quad cell surface 16a relative to the line of sitereference point of the x'-y' coordinate system defined by the quad cell16. Also, based on the signals output by the quad cell 16, the CEDB 17performs a computation to determine the amount of offset (e.g., theangular displacement) between the location of the centroid of star lightenergy distribution 21 on the quad cell surface 16a and the line ofsight reference point (i.e., the intersection of axes x' and y' of quadcell 16). These computations are collectively represented by block (B)of FIG. 5a, and may be performed using conventional techniques. As waspreviously described, the amount of offset (e.g., angular displacement)between the location of the centroid of star light energy 21 on the quadcell surface 16a and the line of sight reference point of the quad cell16 corresponds to the angular displacement (θ₁) of the spacecraft 10.

Information representing the determined position of the centroid of starlight energy distribution 21 on the quad cell surface 16a and the offsetof this distribution 21 relative to the line of sight reference point isoutput from the CEDB 17 to the controller 1. For the purposes of thisdescription, this information is referred to as an error signal (e1),which is represented in FIG. 8. Also, block 24 of FIG. 8 represents thecomputations performed by the CEDB 17 described above.

Within the controller 1, based on the portion of error signal (e1)representing the determined offset of the centroid of star light energy21 relative to the line of sight reference point, a determination ismade of whether the determined offset is equal to or exceeds apredetermined threshold. Preferably, the predetermined threshold is onethat allows the orientation of the telescope 9, LOS, and FOV relative tothe star 22 to change at least somewhat from the desired attitude ofthese elements, and there is an error margin between the predeterminedthreshold and an offset where the orientation of the telescope 9 and FOVrelative to star 22 would be such that the star 22 would be outside ofthe FOV of the telescope 9.

For the purposes of this description, it is assumed that based on theerror signal (e1) representing the position and offset of the centroidof star light energy distribution 21 on the quad cell surface 16arelative to the line of sight reference point at the second instance oftime, the controller 1 determines that the determined offset does notequal or exceed the predetermined threshold. As such, no controlling ofthe telescope 9 is necessary in this case.

A determination that the determined offset is equal to or greater thanthe predetermined threshold indicates that the spacecraft attitude hasfurther varied from the desired spacecraft attitude, owing to, e.g., theenvironmental forces impinging on the spacecraft 10. For example, it isassumed for the purposes of this description that after the occurrenceof the second instance of time, the spacecraft 10 continues to rotateabout the yaw axis as a result of the environmental forces, and that,eventually, at a third instance of time, the spacecraft 10 reaches anangular displacement of (θ₂) (about the yaw axis). This is representedin FIG. 3c. As can be appreciated, if the telescope 9 were stabilizedduring this continued rotation of the spacecraft 10, the orientation ofthe telescope (and the LOS and FOV) relative to the star 22 wouldcontinue to vary accordingly at the same angular velocity as that of thespacecraft 10. However, in accordance with this invention, thecomponents 1 and 5 operate in conjunction with one another to maneuverthe telescope 9 during the continued rotation of the spacecraft 10 so asto cause the telescope 9, LOS and FOV to maintain acceptableorientations relative to the star 22, as will now be described.

As was described above, it is assumed that at the third instance oftime, the spacecraft 10 reaches an angular displacement of (θ₂) (aboutthe yaw axis). It is also assumed that at this instance of time thecontroller 1 controls the quad cell 16 so as to cause signals to be readout from the quad cell 16 to the CEDB 17, and that, as a result, theCEDB 17 outputs to the controller 1 an error signal (e1) representingthe position and offset of the centroid of star light energydistribution 21 on the quad cell surface 16a relative to the line ofsight reference point.

In response to receiving this error signal (e1), the controller 1 makesa determination of whether the offset represented by the signal (e1) isequal to or exceeds the assumed to correspond to the rate (i.e., angularvelocity) at which the spacecraft 10 (and hence, the telescope 9)rotates about the z (yaw) axis as a result of the environmental force(s)impinging on the spacecraft 10.

After the controller 1 performs the calculations described above todetermine the approximate amount of angular displacement of thespacecraft 10 about the yaw axis, to generate the angular torquecommand, and to determine the approximate orientation of the spacecraft10 relative to the star 22, the controller 1 controls the telescopedrive 5, based on the results of these calculations, so as to cause thetelescope drive 5 to apply torque(s) to the telescope 9 for maneuvering(e.g., rotating) the telescope 9 (block F). In the example describedherein, the controller 1 controls the telescope drive 5 so as to apply atorque to the telescope 9 for maneuvering the telescope 9 in an oppositedirection than the direction of motion of the spacecraft 10 about the zaxis (e.g., the telescope 9 is rotated in an opposite direction aboutaxis z than the direction of rotation of the spacecraft) by a totalangular displacement that is substantially equal to the total angulardisplacement (θ₂) of the spacecraft 10. The torque is applied by thetelescope drive 5 in such a manner that the telescope 9 is caused torotate at an angular velocity that is substantially equal to thespacecraft angular velocity (the angular velocity of the telescope 9 isdetermined by the angular torque command determined previously at blockE). That is, based on values of the spacecraft angular displacement,orientation relative to star 22, and angular torque command calculatedpreviously by the controller 1 at blocks E and F, the controller 1controls the telescope drive 5 so as to apply a torque to the telescope9 for causing the telescope 9 to be rotated at an angular velocity thatis substantially equal to the spacecraft angular velocity, and forcausing the telescope 9 (and the telescope LOS and FOV) to have acorrected predetermined threshold. Assuming that the controller 1determines that the offset does equal or exceed the predeterminedthreshold (block C of FIG. 5a), then the controller 1 performs furthercalculations. In particular, based on the error signal (e1) receivedfrom the CEDB 17, the controller 1 performs a calculation to make anapproximate determination of the amount of angular displacement of thespacecraft 10 about each individual axis x, y, and z. In the exampledescribed herein, it is assumed that the spacecraft rotates about onlythe z (yaw) axis. Thus, in this example it is assumed that thecalculation performed by the controller 1 results in an approximatedetermination of the spacecraft 10 displacement about this z axis. Alsobased on the error signal (e1) received from the CEDB 17, the controller1 performs a calculation to make an approximate determination of theorientation of the spacecraft 10 relative to the star 22. Thesecalculations performed by the controller 1 are collectively representedby block D of FIG. 5a.

Moreover, typically the controller 1 performs a further calculation togenerate an angular torque command using the following equation (EQ1):##EQU1## wherein S represents a Laplace Transform that is assumed toemploy the portion of error signal (e1) representing the offset betweenthe centroid of star light energy distribution 21 and the line of sightreference point, T₁ represents a lead time constant, and T₂ represents alag time constant. The calculation performed by the controller 1 forgenerating the angular torque command is represented by blocks 25 and(F) of FIGS. 8 and 5a, respectively. At least a portion of the generatedangular torque command is orientation relative to the star 22 such thatthe centroid of the distribution of star light energy 21 on the quadcell surface 16a is again located at the line of sight reference pointdefined by the quad cell 16 (see, e.g., FIG. 4c). In this manner, thetelescope 9, LOS and FOV are again oriented in their desiredorientations relative to the star 22, regardless of the undesiredrotation of the spacecraft 10. A representation of a correctedorientation of the telescope 9, LOS and FOV relative to star 22, at atime which is assumed to be the third instance of time, is depicted inFIG. 3c. The step of controlling the telescope 9 is represented by block30 in FIG. 8, and, as can be appreciated in view of FIG. 8, the aboveoperations for controlling the position of the telescope 9 are performedin a closed-loop manner based on outputs of the quad cell 16.

In accordance with an aspect of this invention, the SIARSCS 1' alsooperates as a closed-loop control system for providing further controlof the angular velocity of the telescope 9 about the individual axes z"and x", to ensure that the angular velocity of the telescope 9 issimilar in magnitude to that of the spacecraft 10 during the undesiredmotion of the spacecraft 10. By example, as the telescope 9 is beingmaneuvered in the above-described manner, the rate sensor/encoder block3b detects the angular velocity of the telescope 9 about each individualaxis x" and z", and outputs information representing the detectedangular velocity about each individual axis x" and z" to the controller1 (block E1 of FIG. 5b). In the example described above, the telescope 9is rotated about axis z" only, so it is assumed that the informationprovided by the block 3b to the controller 1 represents the detectedangular velocity of the telescope 9 about this axis z" only. Within thecontroller 1, the information representing the detected angular velocityof telescope 9 is compared to a value of the angular torque commandrepresenting the approximate angular velocity of the spacecraft 10(calculated previously at block E) (block 26 of FIG. 8) For cases inwhich the value of the detected telescope angular velocity differs fromthe spacecraft angular velocity rate value represented by the angulartorque command, an error signal (e2) (FIG. 8) is generated which isproportional to the difference between these values. This step isrepresented by block E2 of FIG. 5b. The controller 1 then employs theerror signal (e2) to control the telescope drive 5 for causing the drive5 vary the angular velocity of the telescope 9 by an amount that issubstantially equivalent to the difference represented by the errorsignal (e2) (see, e.g., blocks F3 and 30 of FIGS. 5b and 8,respectively). In this manner, the difference between the calculatedangular torque command (representing a calculated value of thespacecraft angular velocity) and the detected angular velocity isminimized so that the telescope 9 is rotated at substantially the sameangular velocity as that of the spacecraft 10, for maintaining thetelescope 9 in its desired orientation relative to star 22.

Another aspect of the invention will now be described. As was previouslydescribed, the position sensor/encoder block 3a detects the angularposition of the telescope 9 relative to each axis z" and x" and providescorresponding information representing the detected angular positionrelative to each axis z" and x" to the controller 1. In a preferredembodiment of the invention, the position sensor/encoder block 3a has acapability for detecting angular displacements of the telescope 9 (aboutthe individual axes z" and x") that are at least as small as 10 arcseconds, and provides an output representing the angular position of thetelescope 9 about each respective axis z" and x" for each time it isdetected that the telescope 9 has rotated about the individual axis byan increment of approximately 10 arc seconds. Also in a preferredembodiment of the invention, for each of these 10 arc second incrementsabout individual ones of axes z" and x", the rate sensor/encoder block3b provides information to the controller 1 representing the detectedangular velocity of the telescope 9 about the individual axis z" x".

In accordance with the invention, the controller 1 performs aninterpolation to estimate the position of the telescope 9 during timesbetween those at which angular position detections are made by the block3a. By example, FIG. 2a shows a representation of exemplary telescopeangular position values, as detected by the block 3a at times T2, T4,and T6, and FIG. 2b shows a representation of exemplary angular velocityvalues of the telescope 9, as detected by the block 3b at times T2, T4,and T6. In accordance with this aspect of this invention, the controller1 employs the values received from the blocks 3a and 3b to interpolateangular position values for the telescope 9 in between exemplary timesT2, T4, and T6, using the following equation (EQ2):

    θ(T)=θ(t)+θ(t)(T-t)(EQ2)                 (EQ2)

where θ(T) represents the angular position of the telescope 9 at aselected time T, θ(t) represents the angular position of the telescope 9at a particular previous time (t)(e.g., T2, T4, or T6), θ(t) representsthe angular velocity of the telescope 9 at this previous time (t), and(T-(t)) represents a temporal difference between the times T and (t). Anexample of telescope position values (also referred to as inertialposition values) calculated by the controller 1 at selected times (T) ofT3, T5, and T7 is shown in FIG. 2c, superimposed over the exemplaryvalues of FIG. 2a. The interpolation performed by the controller 1 isalso represented by block 7 of FIG. 8.

It should be noted that the times (T) may be determined in accordancewith desired sampling rates (e.g., 1/50 seconds) of the control loop,which sampling rates may be selected based on the brightness of stellarbody 22 and the size of telescope 9.

A further aspect of the invention will now be described. As waspreviously described, for cases in which it is necessary to correct theorientation of the telescope 9 and FOV relative to the star 22 as aresult of undesired spacecraft motions, the SIARSCS 1' operates tomaneuver the telescope 9 so as to maintain the telescope 9, LOS and FOVin a desired orientation relative to the star 22. By example, for theexample described above the telescope 9 is maneuvered in an oppositedirection than that of the spacecraft 10 motion, by an angulardisplacement that is substantially equal in magnitude to the totaldisplacement of the spacecraft 10 about the individual axis z. Also inthis example the telescope 9 is maneuvered about axis z" at a respectiveangular velocity that is substantially equal in magnitude to the angularvelocity of the spacecraft 10 about individual axis z. This operationenables the telescope 9 (LOS, and FOV) to have an orientation relativeto the star 22 such that the centroid of the distribution of star lightenergy 21 on the quad cell surface 16a is maintained at the line ofsight reference point defined by the quad cell 16, even though thespacecraft 10 is in motion. Being that the telescope 9 is maneuvered bya similar angular displacement and velocity as those of spacecraft 10,but in an opposite direction, it can be appreciated that while the starlight energy 21 is maintained at the line of sight reference pointdefined by the quad cell 16, the magnitude of the telescope's angularvelocity detected by the block 3b represents the magnitude of thespacecraft's inertial angular velocity rate as well as that of thetelescope 9, and the angular position of the telescope 9 (as detected byblock 3a and interpolated in the controller 1) corresponds to thespacecraft inertial position. This being the case, it can also beappreciated that the attitude (position) of the spacecraft 10 may bedetermined and controlled based on the detected angular velocity anddetected/interpolated angular position of the telescope 9. As such, andin accordance with this invention, the values of the telescope's angularvelocity and angular position (including the interpolated values) abouteach axis x" and z" are output by the controller 1 to the spacecraftattitude determination and control system (SADACS) 8. The SADACS 8employs these values to determine the attitude of the spacecraft 10 andto control the attitude adjustment actuators 11, if needed, so as toprovide appropriate roll, pitch, and/or yaw torques for controlling thespacecraft attitude for maintaining the attitude of the spacecraftwithin pre-established deadband limits. It should be noted that forcases where the spacecraft attitude is controlled in this manner by theSADACS 8 and attitude adjustment actuators 11, the SIARSCS 1' operatesin a closed-loop fashion, in a similar manner as was described above,for maintaining the orientation of the telescope 9 in its desiredorientation relative to star 22.

As can be appreciated in view of the foregoing description, the SIARSCS1' enables fine attitude control of the spacecraft 10 to be provided,without needing conventional elements, such as gyros or conventionalstellar body trackers, which are typically employed in conventionalspacecraft for generating spacecraft inertial rate and positioninformation, respectively. As such, no gyros or conventional stellarbody tracking equipment need be employed, and, as a result, the costsassociated with these items can be avoided, although such items may beemployed if desired to further enhance spacecraft attitude control.

In accordance with a further aspect of the invention, various,radiation-sensitive ones of the components of the SIARSCS 1' are mountedto the spacecraft 10 at a location which enables the components to beshielded from solar radiation 12 impinging on the spacecraft 10 by atleast a portion of the spacecraft structure, so that the solar radiation12 is prevented from coming into contact with these components. As anexample, FIG. 6 shows the spacecraft 10 and the telescope 9, and alsoshows a block 26, which represents collective components 14, 15, 16, 1,8, 3a, and 3b of the SIARSCS 1'. In accordance with the illustratedembodiment, the block 26 is mounted to a lower surface 10' of a portion10a of the spacecraft 10. At least this portion 10a of the spacecraft 10protects the block 26 by preventing the solar radiation 12 impinging inthe spacecraft 10 from coming into contact with the block 26. As aresult, the operation of the components 14, 15, 16, 1, 8, 3a, and 3b isnot detrimentally affected by the solar radiation 12 impinging on thespacecraft 10, and the SIARSCS 1' can be considered to be "radiationhardened".

It should be noted that the invention is not limited to employing atelescope such as the telescope 9 shown in FIG. 1 and described above.By example, and in accordance with an alternate embodiment of theinvention, a SIARSCS 1" may be provided that includes similar componentsas the SIARSCS 1' of FIG. 1, except that the SIARSCS 1" includes atelescope 32 having a gimballed mirror 13", as shown in FIG. 7. Thetelescope 32 includes similar components as the telescope 9 describedabove, except that mirror 13" is supported by a gimbal support 13'(which is fixed to the telescope structure) which allows the mirror 13"to be rotated about an axis x'". Also, the SIARSCS 1" further comprisesa mirror drive mechanism 34, a position sensor/encoder block 33a and arate sensor/encoder block 33b.

Also in this embodiment, bearings 18b are disposed between a surface 32aof the telescope 32 and a structural portion 19' of a spacecraft 10',and provide gimbal supports for the telescope 32 for allowing thetelescope 32 to be rotated about axis z". A telescope drive mechanism 5'is drivably-engaged with the surface 32a of the telescope 32, and iscontrollable by the controller 1 for being actuated to provide selectedtorques to the telescope 32 for causing the telescope 32 to rotate aboutthe axis z" by a selected angular displacement relative to thespacecraft structure 19', preferably without affecting the attitude ofthe spacecraft 10.

The mirror drive mechanism 34 is drivably engaged with the mirror 13",and is controllable by the controller 1 for exerting a torque whichcauses the mirror 13" to rotate about axis x'" by a correspondingangular displacement, without affecting the attitude of the telescope32. Preferably, the telescope drive mechanism 34 and gimbal support 13'have a collective capability for enabling the mirror 13" to be rotatedabout the axis x'" by at least ±4 degrees.

The position sensor/encoder block 33a detects the angular position(i.e., the arc of rotation) of the mirror 13" relative to axis x'", andtranslates the detected angular position to corresponding informationwhich is provided to the controller 1. The rate sensor/encoder block 33bdetects the angular velocity at which the mirror 13" rotates about theaxis x'", and translates the detected angular velocity to correspondinginformation which is also provided to the controller 1.

Also, a position sensor/encoder block 3a' detects the angular position(i.e., the arc of rotation) of the telescope 32 relative to the axis z",and translates the detected angular position to correspondinginformation which is provided to the controller 1. A rate sensor/encoderblock 3b' detects the angular velocity at which the telescope 32 rotatesabout the axis z", and translates the detected angular velocity tocorresponding information which is also provided to the controller 1.

In accordance with this embodiment of the invention, the SIARSCS 1'operates in a similar manner as was described above, except that inappropriate cases where it is necessary to compensate for a rotation ofthe spacecraft 10' about axis x, the mirror 13", instead of thetelescope 32, is controlled for being rotated about the axis x'" by aselected angular displacement. Also, outputs of the blocks 33a and 33b,in addition to those of the blocks 3a' and 3b', are employed forenabling spacecraft inertial rate and position information to bedetermined, for subsequent use in controlling the attitude of spacecraft10'. That is, spacecraft inertial rate and position information isderived based on information representing the detected position of thetelescope 32, information representing the detected rate at which thetelescope 32 is maneuvered, information representing the detectedposition of the mirror 13", and information representing the detectedrate at which the mirror 13" is maneuvered.

Having described a number of embodiments of the invention, some of theadvantages provided by the SIARSCS 1', 1" of the invention will now bedescribed. By example, and as was previously described, one advantage ofthe SIARSCS 1', 1" is that it is less expensive than conventionalspacecraft attitude control systems that employ gyros and conventionalstellar body trackers. Another advantage is that the SIARSCS 1', 1" is athermally stable system, and the telescopes 9 and 32 each provide a highdegree of detection accuracy as compared to conventional telescopeshaving comparably-sized apertures. Moreover, and as was also previouslydescribed, the SIARSCS 1', 1" enables both satellite-related inertialposition and inertial rate information to be generated, and thus doesnot require the use of conventional stellar body trackers and gyros forgenerating this information. Furthermore, being that the field of view(FOV) of the telescopes 9 and 32 is preferably about ±0.5 degrees, aquad cell detector 16 may be employed which has a smaller area thanradiation detectors employed in conventional telescopes.

It should be noted that any suitable type of drive mechanism may beemployed for the telescope drive 5, including, by example, anelectro-mechanical or inertial drive mechanism.

In the preferred embodiment of the invention, the telescope 9 comprisesa mirror 13, a primary lens 14, a secondary lens 15, a quad celldetector (also referred to as a quad cell) 16, and a centroid errordetection block (CEDB) 17, which operates in a manner as will bedescribed below. Wavelengths 2 (such as those emitted by a stellar bodywithin the field of view (FOV) of telescope 9) received by the telescope9 through an aperture 9b of the telescope 9 are reflected from themirror 13 to the primary lens 14, which then focusses the wavelengths itreceives to the secondary lens 15. Those ones of the wavelengths thatare reflected to the secondary lens 15 are focussed by the secondarylens 15 to a radiation sensitive surface 16a

As was previously described, it should be noted that although theinvention is described above in the context of employing only the blocks3a, 3b (or 33a and 33b) to generate inertial position and rateinformation, gyros and/or stellar body trackers may also be employed inconjunction with the blocks 3a and 3b (or 33a and 33b) to provide suchinformation. Also, in one embodiment, gyros may be employed forgenerating the inertial rate information, and the SIARSCS 1', 1" may beemployed to provide this information in cases wherein one or more of thegyros fail(s) to operate effectively.

While the invention has been particularly shown and described withrespect to preferred embodiments thereof, it will be understood by thoseskilled in the art that changes in form and details may be made thereinwithout departing from the scope and spirit of the invention. Byexample, it should be noted that it is within the scope of thisinvention for the computations performed within the CEDB 17 to beperformed within the controller 1. Also, other forms of equations thatthose described above may be employed for the star tracking procedure.

What is claimed is:
 1. In a satellite having a star tracking system,said tracking system including an optical sensor mounted on thesatellite for relative movement thereon, a light sensitive panel mountedto receive light from the light sensor and convert said light to anelectrical signal, a processor for determining the position of the lightsignal on the panel, and a drive to move the optical sensor, a method ofcontrolling the optical sensor comprising the steps of:aligning saidoptical sensor with a star that provides a spatial reference for thesatellite; focusing light from said star on the light sensitive panel ata predetermined panel reference location; sensing the position of thefocused light on said panel over a period of time; comparing saidreference location with the position of said focused light; generatingan error signal representing the displacement of said focused light fromsaid reference location; actuating said drive in response to said errorsignal to move said optical sensor relative to the satellite to returnthe focused light to the reference location; and sensing the positionand rate of movement of the optical sensor relative to the satellite andprocessing said sensed data to derive the inertial rate and position ofthe satellite.
 2. In a satellite having a star tracking system, saidtracking system including an optical sensor mounted on the satellite forrelative movement thereon, a light sensitive panel mounted to receivelight from the light sensor and convert said light to an electricalsignal, a processor for determining the position of the light signal onthe panel, and a drive to move the optical sensor, a method ofcontrolling the optical sensor, as described in claim 1, wherein thestep of actuating said drive occurs only when said displacement is equalto or greater than a predetermined minimum threshold.
 3. In a satellitehaving a star tracking system, said tracking system including an opticalsensor mounted on the satellite for relative movement thereon, a lightsensitive panel mounted to receive light from the light sensor andconvert said light to an electrical signal, a processor for determiningthe position of the light signal on the panel, and a drive to move theoptical sensor, a method of controlling the optical sensor, as describedin claim 1, further comprising the step of adjusting the attitude of thesatellite in response to said derived inertial rate and position of thesatellite.
 4. In a satellite having a star tracking system, saidtracking system including an optical sensor mounted on the satellite forrelative movement thereon, a light sensitive panel mounted to receivelight from the light sensor and convert said light to an electricalsignal, a processor for determining the position of the light signal onthe panel, and a drive to move the optical sensor, a method ofcontrolling the optical sensor, as described in claim 1, wherein saidsteps are performed continuously in response to changes in attitude ofthe satellite.
 5. In a satellite having a star tracking system, saidtracking system including an optical sensor mounted on the satellite forrelative movement thereon, a light sensitive panel mounted to receivelight from the light sensor and convert said light to an electricalsignal, a processor for determining the position of the light signal onthe panel, and a drive to move the optical sensor, a method ofcontrolling the optical sensor, as described in claim 1, wherein saiddrive is actuated to move the optical sensor at an angular ratesubstantially equal to the angular rate of the movement of thesatellite.
 6. In a satellite having a star tracking system, saidtracking system including an optical sensor mounted on the satellite forrelative movement thereon, a light sensitive panel mounted to receivelight from the light sensor and convert said light to an electricalsignal, a processor for determining the position of the light signal onthe panel, and a drive to move the optical sensor, a method ofcontrolling the optical sensor, as described in claim 1, furtherincluding the step of mounting the optical sensor on the satellite in aposition such that a portion of the satellite shields the optical sensorfrom solar radiation.
 7. A star tracking system for an orbitingsatellite comprising:an optical sensor mounted on the satellite forrelative movement thereon, said optical sensor being aligned with a starthat provides a spatial reference for the satellite, and focusing lightfrom said star on a light sensitive panel, said optical sensor furtherincluding a drive to move said sensor relative to said satellite; alight sensitive panel mounted to receive light from the light sensor andconvert said light to an electrical signal, said panel having apredetermined reference location; a first processor for determining theposition of the light on the panel, calculating the displacement of saidposition from said predetermined reference location, and generating anerror signal representing said displacement; a second processor foractuating said drive in response to said error signal to move saidoptical sensor relative to the satellite to return the focused light tothe reference location; and wherein said second processor further sensesthe position and rate of movement of the optical sensor relative to thesatellite and processes said sensed data to derive the inertial rate andposition of the satellite.
 8. A star tracking system for an orbitingsatellite, as described in claim 7, wherein said drive is actuated onlywhen said displacement is equal to or greater than a predeterminedminimum threshold.
 9. A star tracking system for an orbiting satellite,as described in claim 7, wherein said second processor further adjuststhe attitude of the satellite in response to said derived inertial rateand position of the satellite.
 10. A star tracking system for anorbiting satellite, as described in claim 7, wherein said star trackingsystem operates continuously in response to changes in attitude of thesatellite.
 11. A star tracking system for an orbiting satellite, asdescribed in claim 7, wherein said drive is actuated to move the opticalsensor at an angular rate substantially equal to the angular rate of themovement of the satellite.
 12. A star tracking system for an orbitingsatellite, as described in claim 7, wherein said optical sensor ismounted on the satellite in a position such that a portion of thesatellite shields the optical sensor from solar radiation.